This invention relates to a rotor blade, and, in particular to a composite aerodynamic rotor blade assembly, such as a helicopter rotor blade assembly in which the spar heel is separately fabricated and assembled with the aft fairing structure, the aft fairing skin members are secured between the spar heel and the cap member to insure a fail-safe design, and the deicing blanket structural material and aft fairing skin members serve as load carrying members, thereby providing the existing state-of-the-art with a low cost and minimum operation production oriented blade with superior structural integrity.
From the very advent of the composite rotor blade, those skilled in the art have sought to reduce the cost of manufacture by in some way reducing or changing the various stages of fabrication. Still, however, composite rotor blades are fabricated by joining a number of separately fabricated subassemblies; and, for the most part, as multicured subassemblies requiring separate bonding assembly jigs for each major cure subassembly.
For example, in a known method of fabricating a composite rotor blade, the following subassemblies are produced:
1. blade cap member, deicing blanket and nose block subassembly;
2. blade spar subassembly;
3. blade spar and cap member, deicing blanket and nose block subassembly;
4. blade trailing edge wedge subassembly;
5. blade aft fairing core, (unmachined) with one skin member subassembly;
6. blade aft fairing core (machined) with both skin members subassembly;
7. Final assembly including subassemblies 3-6.
As can readily be seen, this assembly includes at least seven curing and/or bonding operations. The fabrication of a blade with this number of curing and/or bonding operations is necessarily costly and less than desirable from this standpoint alone.
It would, therefore, be desirable to be able to reduce the total number of curing and/or bonding operations now required to fabricate a composite rotor blade and thereby reduce the cost of fabrication, while at the same time not adversely affecting the structural integrity of the blade.
Of the various subassemblies mentioned above, one of the most limiting to the achievement of production economy and optimized structural integrity is the spar subassembly. For example, in one known method of fabrication which employs curing, it has been found that a back pressure has to be provided against the rear face of the spar to counteract the internal bag pressure acting within the spar during the curing cycle to avoid possible structural damage. In the past the only successful way to do this was by forming the spar separately in a mold. In another known method of fabrication which employs curing, it was decided to fabricate the spar heel separately from the spar and to then include the spar heel in assembly with the spar during the spar curing cycle. This procedure, however, did not prove satisfactory because, for one thing, the spar developed undesirable surface wrinkles which hampered surface bonding and consequently load transfer to other parts of the blade.
It would, therefore, be desirable to provide a composite rotor blade according to which the spar design assembly is improved and does not develop any undesirable conditions detrimental to the proper employment of the spar.